1. Technical Field of the Invention
The present invention is generally directed to structural vibrations, and particularly to a blading system and method for controlling structural vibrations in axial-flow compressors, turbines or fans, as in aircraft engines and like turbomachines. This invention more specifically relates to a damper which is applied externally to the dovetail root attachments of rotor blades, and which does not necessitate structural modification to the rotor blades or the turbomachine incorporating these blades. The damper significantly suppresses destructive structural vibration, without affecting the aerodynamic performance of the turbomachine.
2. Description of the prior Art
The reliability of an axial-flow turbomachine in general, and a compressor in particular, depends on the durability of its component parts. The blading system of the turbomachine can, in the event of vibration fatigue, dangerously jeopardize such durability. Moreover, the geometry of the compressor blading system may enhance the vibration susceptibility. Vibration failures have generally been attributed to several sources, among which are the followings:
1. Forced vibrations due to mechanical obstructions which cause unequal pressure distributions in the blade path; PA1 2. aerodynamic self-excited vibrations due to flutter phenomena; and PA1 3. vibrations due to the rotating stall. PA1 1. Internal friction of the material; PA1 2. dissipation of vibration energy to the environment; PA1 3. transfer of vibration energy through the blade root and rotor to other members in the system having similar natural frequencies; and PA1 4. dissipation of vibration energy by the mounting.
Considerable knowledge has been gained in recent years on the mechanism of vibration; however, the complexities introduced by combinations of the sources of vibration continue to make vibration fatigue an important problem in turbomachine operation. Many attempts to reduce vibration fatigue are based on the premise that this problem cannot always be thoroughly considered in the blading design, and therefore must be dealt with in the development or operation of the turbomachine. These solutions could require extensive structural and design modifications to the turbomachine as well as to the production and quality control lines.
Since the principal causes of vibration continue to be ever present, the possible suppression of the vibration lies in the amount of damping available to the blading system. It has been generally accepted, so far, that the inherent damping of the conventional blading system is insufficient in some instances to prevent blade failure. Reference is made to "A Vibration Damper for Axial-Flow Compressor Blading", M. P. Hanson, National Advisory Committee for Aeronautics, S.E.S.A. Proceedings, Volume XIV, No. 1.
Vibration induced fatigue failure of rotor blades is of a continuing concern to designers who take into account the following blade design criteria: Blade life prediction, allowable frequency margins which need to be justified on the basis of the intensity of resonant stresses, and accurate prediction of susceptibility to aeroelastic instabilities.
Contributions to inherent damping in a vibrating rotor blade stem from four principal sources:
It has been believed that material damping is limited in currently used materials and has proven insufficient in systems susceptible to high amplitude vibrations. Aerodynamic damping, while very effective in reducing the vibration resulting from certain types of excitation, cannot be relied upon to dampen flutter types of vibration in which the very source of excitation is aerodynamic instability. The vibration-absorption method of providing damping is limited to systems having several components of similar natural frequencies. In compressors, it is difficult to maintain rotor blades at desired frequency ratios because of manufacturing tolerances, variations in mounting, and frequency changes due to centrifugal forces. Vibration energy dissipation in the mounting is due to relative motion between the blade and the rotor in mechanical blade attachments. At low rotor speeds the energy dissipation is appreciable; however, at high rotor speeds the centrifugal force essentially tightens the blade sufficiently to destroy the dissipation.
Damping, in the context of this application, refers to energy dissipation capacity of the component undergoing vibration. A measure of energy dissipation is defined as the ratio of energy lost per cycle of vibration to maximum vibration energy stored in the appropriate mode divided by 2. Where such a measurement is inadequate or impractical, the extent of damping present is measured indirectly as attenuation of strain in the vibrating components due to the influence of the damping mechanism.
Material or hysteretic damping refers to energy dissipation due to many complex mechanisms within a material, when a volume of the material is subjected to cyclic stresses. Accurate measurements show that cyclic stress-strain always has a hysteresis loop. Thus, material damping is always present in a vibrating rotor blade.
Friction damping at an interface depends on loading, roughness of surfaces, level of external excitation, slip amplitude, geometry of the contacting components, etc. For example, the root structure of a jet engine blade is typically of a dovetail, pin or fir tree design, the extent of friction damping being different in each design. In the case of shrouded blades, untwisting of the blades under centrifugal loading brings neighboring blades into contact at shrouds, resulting in a complex "joint" at the interface. Reference is made to "Turbojet Engine Blade Damping", A. V. Srinivasan et al., NASA Contract Report 165406, July 1981.
So far, no single analytical or experimental approach appeared feasible for the evaluation of all sources of damping, due to the diversity of mechanisms involved. Several methods have been proposed to limit vibration susceptibility in blade mechanisms. One of the most obvious methods is to control the source of vibration, such as by avoidance, which can prove to be inefficient and quite difficult to achieve. For instance, in some applications, such as in wind tunnels where axial-flow compressors are employed, avoidance may be achieved by reducing the energy level of the system in order to reduce the rotor blade vibrations. However, such a reduction in the energy level causes the system to operate below its intended operation design and optimal efficiency.
Another common method to limit vibration has been to increase the blade stiffness by increasing the blade thickness, shrouding the blades, or lacing the blades with wires. Still an alternative proposed solution has been to improve the material strength by using material that is less sensitive to metal fatigue, such as high fatigue strength metallic alloy or composites. However, these corrective measures may be harmful to the aerodynamic efficiency of the blading system, may create a structural problem in high-speed turbomachines, or may increase the manufacturing cost.
Another conventional method is via mechanical damping, which consists of various means of increasing the damping of a system by dissipating the vibration energy as frictional heat or through another system of similar natural frequencies. Many blade damping devices have been based on these principles, but the design usually adds to the complexity of the manufacture of the blade and sometimes can interfere with its aerodynamic performance. The above article by Hanson purports to address this problem by adding pins that contact the blade root and the rotor. The proposed damping mechanism introduces frictional forces into a compressor blading system, and requires an addition of an appendage at the base of the blade. However, such design modifications cannot be introduced universally into conventional turbomachines, and can prove to be quite expensive and impractical to retrofit into existing blading systems.
The two common types of mechanical dampers are tuned-mass and friction dampers. Tuned-mass damper works on the principle of adding a secondary mass, either internally or externally to the main structure, to absorb the vibration energy at a specific frequency to which the secondary mass is tuned to. An example of tuned-mass damping is the helicopter rotor systems. Friction damper provides a broader frequency range of application and relies on interface rubbing motion to dissipate the vibration energy. Snubber in turbine blade design, which is usually placed at mid-span to form an interlocking ring, is a form of friction damper as well as a stiffness modifier.
Most of these conventional damping methods often require structural modifications to the rotor blades. Such added modifications can reduce the aerodynamic performance of the turbomachine. This is generally the case with the snubber damper design. Tuned-mass damper usually has a narrow frequency range of application, while friction damper will eventually wear out the blades, thereby requiring extensive maintenance. These damping methods entail a high level of complexity and considerable manufacturing, design, and maintenance costs.
Other damping mechanisms of general interest to the present invention have also been designed and patented, some of which are described in the following patents:
U.S. Pat. No. 4,188,171 (1980), to Baskin, titled "Rotor Blade Internal Damper", describes a load absorbing elastic member which is incorporated internally into the design of a rotor blade to assist in reducing flapwise and chordwise bending moments, torsional loading, and to reduce the control system loads induced by rotor blade moment stall. The load absorbing elastic member has a constraining member which serves as a strain amplifier. However, this damping device and more particularly the constraining member, requires extensive and expensive modification to incorporate the device into the internal structure of the blade, and it applicability may be limited to new blade construction made of composite materials. This method cannot be used readily to retrofit the internal dampers into existing rotor blades due to its integral design feature.
U.S. Pat. No. 4,192,633 (1980), to Herzner, titled "Counterweighted Blade Damper", describes a segmented damper provided with a lower dovetail portion and insertable into a dovetail slot for radial retention. The damper is engageable on its face by a blade retainer for axial retention, such that the damper is pivotable in an axial plane. The damper upper portion has on its one side a flange for engaging the blade shank rail and on its other side a flange for shifting the damper center of gravity axially outward from the lower pivot point. Centrifugal force, occasioned by rotation of the disc and damper then causes the damper to pivot and impart an axial force against the blade to thereby dampen vibrations. This method exemplifies the complexity of a typical, conventional mechanical damper.
Therefore, there is still a significant and unsatisfied need for a damper and method for controlling structural vibrations in axial-flow compressors, turbines, fans as in aircraft engines, and like turbomachines. The damper should be applied externally to the various types of root attachments of rotor blades, such as dovetail, fir tree, or other attachments of different designs. The damper should not necessitate structural modification to the blades or the turbomachine incorporating these blades. It should provide significant vibration damping to suppress destructive structural vibration and thereby to increase the durability of the blading system and the turbomachine. Along with the minimization of structural fatigue, the damper should reduce acoustic noise accompanying high amplitude vibrations.